The present invention relates to a miniature gas turbine engine, and more particularly to a combustor system in which each combustion region approximates a well-stirred reactor.
Miniature turbojet engines (100 lbf thrust and smaller) are often utilized in single usage applications such as reconnaissance drones, cruise missiles, decoy and other weapon applications, including air-launched and ground-launched weapon systems. The use of such an engine greatly extends the range of the weapon in comparison to the more conventional solid fuel rocket engine. Miniature gas turbine engines are difficult to fabricate economically for general expendable usage in large numbers.
To achieve economically feasible extended range expendable propulsion sources for such weapon system, it is necessary that the gas turbine engines be manufactured relatively inexpensively yet provide a high degree of reliability and efficiency. One component that greatly affects performance yet is rather complicated to manufacture is the combustor system.
Miniature gas turbine engines typically utilize annular combustor shapes that wrap around other engine features such as an exhaust tailpipe or a turbine wheel to minimize frontal area in order to maximize the thrust per unit drag. If the engine frontal area is minimized, the combustor internal volume must be utilized optimally.
Miniature gas turbine engine combustor systems may not have room for conventional fuel injection systems and require high-density, high-viscosity fuels to maximize thrust. The combustor system must accommodate these fuels and provide reliable ignition and stable operation. These requirements are a challenge given the size and cost limitations for an expendable system.
Accordingly, it is desirable to provide an inexpensive and reliable combustor system having a minimal frontal area for a miniature gas turbine engine which achieves stability throughout the flight envelope and combustion efficiency at cruise conditions.